1. Field of the Invention
Liquid propellants have been successfully used for many years in various types of rocket propulsion systems for many different types of spacecraft. Those with skill in the art will, of course, appreciate that, for one thing, one major advantage of liquid propulsion systems is that the control of the propellants is easily carried out either by feeding the propellants to the rocket engine from a pressurized tank or by utilizing turbine-driven pumps and control orifices in the flow lines to regulate the flow of the propellants into the combustion chamber of the engine.
For various technical reasons it has not been considered advantageous heretofore to utilize gaseous propellants such as oxygen and hydrogen in rocket propellant systems. For one thing, it will, of course, be appreciated that the real-time control of gaseous propellants being used to power a rocket engine is not an easy task even in those situations where the engine is being fired continuously. Moreover, it is recognized that real-time control of gaseous propellants is even further complicated where the rocket engine is being intermittently operated. As a result, until recently there has been little interest in developing methods and apparatus for effectively controlling gaseous propellants for powering spacecraft rocket engines. Recently, however, advances in the spacecraft art have made it worthwhile to give serious consideration toward utilizing reaction control systems that decompose water by electrolysis and then utilize the oxygen and hydrogen that are produced by the process as propellants for a spacecraft rocket engine.
Accordingly, this invention relates to new and improved methods and apparatus for providing real-time control of a gaseous-propellant rocket engine until the propellants are expended. More particularly, the present invention is related to methods and electronic control apparatus for accurately balancing the mass quantities of the gaseous propellants supplied to a rocket engine for spacecraft and insuring their optimum utilization regardless of whether the engine is to be fired continuously or is to be sequentially fired so as to carry out a series of successive operations.
2. Background Art
A typical bi-propellant rocket propulsion system for a spacecraft includes individual tanks respectively carrying a combustible fuel component and an oxidizer component which, as described above, have heretofore typically been liquids such as hydrogen and liquid oxygen. To operate the rocket engine, these liquids are fed to the engine either by pumps or by individually pressuring the tanks. In either event, the liquid propellants are withdrawn from their respective supply tanks and mixed just before they are introduced into the combustion chamber of the engine. To achieve optimum performance of the spacecraft, these combustible components are usually mixed in an 8-to-1 ratio (by mass) for achieving optimum performance of the rocket engine.
A typical liquid-propellant system as seen, for example, in U.S. Pat. No. 4,541,238, is arranged for measuring the flow rates of the fuel and the oxidizer being supplied to the rocket engine by turbine-driven pumps. The measurements are cooperatively correlated by way of suitable electronic control instrumentation for periodically determining the ratio of the propellants that are being introduced into the combustion chamber of the rocket engine. By comparing the computed mixture ratio with a predetermined set value, the control instrumentation cooperatively regulates the flows of the liquid propellants so as to hopefully achieve optimum utilization of the propellants during the operation of the rocket engine. The system disclosed in that patent is also cooperatively arranged to adjust the predetermined set value of the mixture ratio should subsequent measurements indicate a need for a more-efficient mixture ratio. Those skilled in the art will, however, appreciate that it is difficult, if not practically impossible, to readily evaluate or change the various control parameters of a liquid-propellant system of this nature during successive multiple-burn operations of the spacecraft engine.
Another liquid-propellant system for spacecraft is shown in U.S. Pat. No. 4,722,183 in which the mixture ratio of the liquid propellants is periodically adjusted in keeping with the amounts of the propellants which are respectively remaining in the propellant storage tanks following one or more previous operations of the rocket engine. To achieve these mixture-ratio adjustments, measurements are made after each operation of the engine to determine the quantities of the fuel and oxidizer left in their respective tanks. The pressure is then adjusted in at least one of the propellant tanks so that the fuel and oxidizer will be subsequently withdrawn from their respective storage tanks in whatever proportions are required for establishing a desired mixture ratio for the fuel and oxidizer which are supplied to the rocket engine when it is next fired. In this manner, the relative proportions of the remaining liquid propellants will be adjusted from time to time as needed for balancing the consumption of the fuel and oxidizer during the rest of the spacecraft operation.
In U.S. Pat. No. 4,777,794 a propellant flow control system for a typical spacecraft rocket engine is shown as utilizing fluid-operated, self-regulating control valves or so-called "mass flow regulators" respectively arranged to function without electrical controls. To accomplish this, these self-regulating valves are respectively provided with enclosed bellow actuators cooperatively arranged to position control elements to compensate for temperature or pressure changes that typically occur as propellants are withdrawn from their respective storage tanks. In operation, these self-regulating control valves cooperate so as to establish essentially-constant mass flow rates of the propellants as they are being supplied to the rocket engine.
It will, of course, be appreciated by those skilled in the art that a control system designed for controlling the supply of liquid propellants to a spacecraft engine is not necessarily suited for controlling gaseous propellants. For instance, with liquid-propellant control systems such as the systems described in the aforementioned patents, no attempt is made to control the pressures in the engine combustion chamber as the rocket engine is operating. Moreover, these control systems utilized for liquid propellants are simply not designed for establishing adjustable set points for the system control elements for accommodating unexpected changes in the operating parameters of the system which may take place from time to time during the operation of the engine. Furthermore, with the prior-art liquid propellant control systems, there is simply no necessity to accommodate large changes in propellant densities which will inherently occur in gaseous propellant systems.